ground-effect characteristics of the tu-144 supersonic transport
TRANSCRIPT
NASA/TM-2003-212035
Ground-Effect Characteristics of the Tu-144 Supersonic Transport Airplane
Robert E. CurryNASA Dryden Flight Research CenterEdwards, California
Lewis R. OwensNASA Langley Research CenterHampton, Virginia
October 2003
EC97-44203-03
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NASA/TM-2003-212035
Ground-Effect Characteristics of the Tu-144 Supersonic Transport Airplane
Robert E. CurryNASA Dryden Flight Research CenterEdwards, California
Lewis R. OwensNASA Langley Research CenterHampton, Virginia
October 2003
National Aeronautics andSpace Administration
Dryden Flight Research CenterEdwards, California 93523-0273
NOTICE
Use of trade names or names of manufacturers in this document does not constitute an official endorsementof such products or manufacturers, either expressed or implied, by the National Aeronautics andSpace Administration.
Available from the following:
NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS)7121 Standard Drive 5285 Port Royal RoadHanover, MD 21076-1320 Springfield, VA 22161-2171(301) 621-0390 (703) 487-4650
ABSTRACT
Ground-effect characteristics of the Tu-144 supersonic transport airplane have been obtained fromflight and ground-based experiments to improve understanding of ground-effect phenomena for this classof vehicle. The flight test program included both dynamic measurements obtained during descendingflight maneuvers and steady-state measurements obtained during level pass maneuvers over the runway.Both dynamic and steady-state wind-tunnel test data have been acquired for a simple planform model ofthe Tu-144 using a developmental model support system in the NASA Langley Research Center14- by 22-ft Subsonic Wind Tunnel. Data from steady-state, full-configuration wind-tunnel tests of theTu-144 are also presented. Results from the experimental methods are compared with results from simplecomputational methods (panel theory). A power law relationship has been shown to effectively fit thevariation of lift with height above ground for all data sets. The combined data sets have been used toevaluate the test techniques and assess the sensitivity of ground effect to various parameters.Configuration details such as the fuselage, landing gear, canards, and engine flows have had little effecton the correlation between the various data sets. No distinct trend has been identified as a function ofeither flightpath angle or rate of descent.
NOMENCLATURE
a
x
acceleration in X body axis,
g
, positive forward
a
z
acceleration in Z body axis,
g
, positive down
b
span, ft
C
D
drag coefficient
C
L
lift coefficient
C
m
pitching moment coefficient about the reference center of gravity (40-percent
MAC
)
C
N
normal force coefficient, positive up
C
X
force coefficient in X body axis, positive forward
C
Z
force coefficient in Z body axis, positive down
DGE dynamic ground effect
DGPS differential global positioning system
h
height of the wing aerodynamic center above the ground, ft
HSCT high-speed civil transport
I
x
,
I
y
,
I
z
moments of inertia about the X, Y, and Z axes, slugs-ft
2
k constant
MAC
mean aerodynamic chord, ft
n1
stage one compressor rate of rotation for engine number 1, percent
NASA National Aeronautics and Space Administration
OGE out-of-ground effect
2
p
roll rate, rad/sec
q
pitch rate, rad/sec
dynamic pressure, lb/ft
2
r
yaw rate, rad/sec
S
wing reference area, ft
2
T
thrust, lb
V
true airspeed, ft/sec
W
gross weight, lb
α
angle of attack, deg
γ
flightpath angle, positive for ascending flight, deg
δ
e
elevon position, deg
incremental change in lift coefficient due to ground effect
θ
pitch attitude, deg
Subscripts
eff
flight conditions simulated through dynamic wind-tunnel testing
ref
adjusted to reference angle of attack and/or elevon position
uncorrected dynamic wind-tunnel data prior to removal of inertial force components
α
partial derivative with respect to angle of attack
δ
e
partial derivative with respect to elevon position
INTRODUCTION
Aerodynamic ground effect is an important aspect in the design of large transport aircraft, because ofthe potential impact on takeoff and landing performance and flying qualities. Simulation, control systemdesign, and performance predictions depend on accurate ground-effect models, which are usuallyobtained through proven analysis, wind-tunnel test techniques, and experience gained from similarconfigurations.
Slender wing configurations that have been proposed for an advanced high-speed civil transport(HSCT) aircraft (ref. 1) might exhibit ground-effect characteristics that are different from those ofconventional subsonic transport aircraft. This issue was studied during initial supersonic transport aircraftdevelopment efforts in the 1960s (refs. 2, 3). Accepted aerodynamic theory indicates that ground effect ata particular nondimensional height above ground (
h/b
) will generally be more significant forcomparatively lower-aspect ratio aircraft. Although ground effect has generally been approached as asteady-state situation, studies indicate that the rate of descent of an aircraft during landing might alsoinduce “dynamic” aspects to the problem. Canards, aft-mounted underwing engines, and the presence of
q
∆CLGE
3
vortex lift are other potential HSCT features that might influence ground-effect characteristics. Theimportance of these issues has emphasized the need for further study of ground effect and the reliabilityof test techniques for this new class of vehicle.
Ground-effect prediction has been a challenging problem for many reasons. Analytical methods havegenerally been based on steady-state flow superposition and engineering methods such as panel codes(refs. 4, 5). These methods have limited ability to incorporate configuration complexity such as high-liftsystem details and the modeling of engine exhaust flows. Wind-tunnel testing often has similarlimitations and might invoke additional complications. A moving ground belt or other devices are oftenneeded to remove the unrealistic boundary layer on the wind-tunnel ground plane simulation. Recentstudies have attempted to also include dynamic effects (refs. 6, 7, 8). References 9–14 discuss ground-testmethods that have been used to address the dynamic aspect of ground effect. Both analytical andwind-tunnel predictive methods have had few opportunities for validation.
Although flight tests might provide data to validate the performance of the predictive methods, thesemeasurements are difficult to obtain. References 15–21 describe ground-effect measurements obtainedfor full-scale vehicles in flight and compare the measurements with predictive data. The parametricvariations are limited when an aircraft is flown in close proximity to the ground. The vehicle must be keptclose to trim, and sink rates must be controlled to avoid overstressing the landing gear. Flightmeasurements can be obtained in both steady-state (level flight over the ground) and dynamic(descending flightpath similar to landing) conditions; however, relating the ground-effect increments to areliable out-of-ground effect (OGE) reference condition can be difficult. Because ground-effectincrements are relatively small compared to other vehicle forces and moments, even small atmosphericdisturbances can affect the quality of the flight data. Any wind across the runway has a boundary layer, orvarying velocity profile, that will provide a systematic error in the ground-effect measurements.
As part of the NASA High-Speed Research program, the Tupolev Design Bureau modified a Tu-144(Tupolev Aircraft Company, Moscow, Russia) supersonic transport aircraft for use as a flying test bed(ref. 22). The configuration and performance capability of the test bed, redesignated the Tu-144LLaircraft, are relevant to many HSCT issues, and seven experiments have been conducted (refs. 23, 24).One of these experiments was directed at the study of ground effect, which is the subject of this report.
The objective of the Tu-144 flight experiment was to measure the ground-effect characteristicsthrough a range of flight conditions including steady-state and dynamic conditions. These results providea database for evaluation of experimental and theoretical predictive methods.
Concurrent to the flight activity, a unique wind-tunnel model support system was developed to enablesimulation of dynamic ground-effect conditions in the NASA Langley Research Center(Hampton, Virginia) 14- by 22-ft Subsonic Wind Tunnel. During initial evaluations of the developmentalsystem, a simple planform model of the Tu-144 airplane was tested. Results from these pathfinder tests,conducted under both dynamic and steady-state conditions, were obtained for correlation with the flightdata.
Steady-state, full-configuration wind-tunnel data for the Tu-144 in ground effect are also included inthis study. These data were obtained during initial development of the Tu-144 airplane.
4
This report describes the flight and wind-tunnel test techniques and data analysis methods.Measurement uncertainties, issues, and limitations of the various methods are discussed. Ground-effectdata from the experimental methods are compared with predictions from engineering methods (paneltheory). A power law relationship for the ground-effect increments with respect to height above ground ispresented and used to fit results from the various data sets for more detailed comparisons. The data setsare used to assess the significance of parameters (such as sink rate and configuration complexity) and tocorrelate with results from other aircraft configurations.
VEHICLE DESCRIPTION
A Tu-144 “D” model airplane, originally configured as a commercial supersonic passenger jet, wasmodified into a test bed for high-speed flight research (ref. 22). As part of the modification process, theoriginal RD-36-51 engines were replaced with NK-321 engines (named for N.D. Kuznetsov, SamaraScientific and Technical Complex, Samara, Russia), and a research instrumentation system was installed.Flight tests were conducted at Zhukovsky Air Development Center (near Moscow, Russia) using themodified Tupolev Tu-144 supersonic transport airplane.
Figure 1 shows a three-view of the Tu-144 airplane, and table 1 summarizes essential physicalproperties. The vehicle has a double-delta-wing planform with an inboard leading-edge sweep angle of76° and an outboard leading-edge sweep angle of 57°. The four engines are mounted in two separate podsbelow the wings. Pitch control is provided by symmetric deflection of the wing trailing-edge surfaces orelevons. Roll control is provided by differential deflection of the elevons.
Figure 1. Three-view of the Tu-144 airplane.
020553
5
During normal landings, the sharp nose section is canted downward (drooped) 17° to improve thepilot’s field of view. For all maneuvers in this study, the nose was drooped and the landing gear weredeployed. A highly cambered, unswept canard can be deployed just aft of the cockpit during normallandings. Maneuvers for this study were conducted with the canard retracted and deployed. Figure 2shows a photograph of the Tu-144LL airplane in the normal landing configuration.
Figure 2. Tu-144LL airplane in the normal landing configuration.
Table 1. Physical properties of the Tu-144 airplane.
Reference area, ft
2
5457.8
Reference chord (
MAC
), ft 76
Reference span, ft 94.8
Reference center of gravity, percent
MAC
40
Inboard leading-edge sweep, deg 76
Outboard leading-edge sweep, deg 57
Aspect ratio 1.65
Length overall, ft 220
Typical weight during test maneuvers, lb 265,000
EC97-44203-03
6
FLIGHT TESTING
This section discusses the flight testing of the Tu-144 airplane. The measurements, derivedparameters, flight maneuvers, and data analysis are described in detail.
Measurements
The airplane was equipped with an extensive instrumentation system to support a variety of flightexperiments including ground-effect research. A detailed description of this system is presented inreference 22. Data from a total of 1178 channels were acquired and encoded by a 12-bit pulse codemodulation (PCM) system. The PCM stream was recorded on board with a digital tape recorder andpersonal computer. Primary measurements for the ground-effect data analysis were body axis linearaccelerations, angular attitudes, angular rates, and control surface positions. Lateral and verticalaccelerations were recorded at approximately 64 samples/sec, and the other flight dynamics parameterswere recorded at 32 samples/sec.
An Ashtech Z-12 (Ashtech, Inc., Sunnyvale, California) carrier phase differential global positioningsystem (DGPS) satellite receiver was installed on board the Tu-144LL airplane. Data were obtained at arate of 2 samples/sec, stored in the internal memory of the unit, and downloaded after each flight. Thesedata were merged with data from a ground-based unit using the Ashtech Precise Differential GPSNavigation and Surveying (PNAV™) software (ref. 25). This software computes time and space positiondata and then derives other useful parameters such as flightpath angle and ground speed. The PNAV™software also determines the root-mean-square uncertainty in the space position data. During the majorityof the ground-effect testing, PNAV™ indicated uncertainties less than 0.3 ft.
Meteorological conditions in the vicinity of the runway were measured with conventionalground-based sensors. Windspeed and direction were recorded for each maneuver.
Derived Parameters
The height of the airplane above the ground was determined from the DGPS data and other onboardmeasurements. This derivation followed the approach described in reference 21 and is believed to offerthe same level of uncertainty (less than ±0.5 ft). The DGPS directly measured the location in space of theonboard global positioning system antenna. The measurement of pitch attitude was used to determine thelocation of other points on the airplane, such as the reference center of gravity and the main landing gear,under the assumption that the airplane is a rigid body. The height above ground of the aerodynamicreference point was determined as the difference between the location of the reference point in space andan analytical model of the runway surface. The accuracy of this method was confirmed by checking thederived height values during times when the airplane was in contact with the runway. Details of theseprocedures and the runway model are presented in reference 22.
The measurement of pressure altitude was not used to determine height above ground, because thesensor calibration could be influenced by ground effect. For the same reason, the conventional onboardmeasurement of angle of attack (from a flow angle vane) was not used for analysis of the ground-effectmaneuvers. Because tests were conducted only during low-windspeed conditions, angle of attack wasderived from the DGPS flightpath angle and pitch attitude.
7
The pitch angle acceleration was determined by differentiating the pitch rate data. Linearaccelerations at the reference center of gravity were determined by adjusting the accelerometer sensorvalues for radius of rotation and angular accelerations.
After each flight, the mass properties were derived from onboard fuel quantities recorded in flight.Thrust levels during the ground-effect maneuvers were estimated postflight using onboard recordedengine parameters.
Flight Maneuvers
Two types of flight maneuvers were used to collect ground-effect data. The first maneuver, called theconstant-alpha approach, is based on a method developed in reference 18. Data were collected as theaircraft descended toward the runway in a “dynamic” ground-effect situation. In the second type ofmaneuver, called the level pass approach, data were obtained as the aircraft flew level passes over therunway in a “steady” ground-effect situation.
Prior to each constant-alpha maneuver, the pilot appropriately configured the airplane by extendingthe landing gear, canting the nose down 17°, and either deploying the canard or leaving it retracted. Theairplane was then aligned in a descent toward the runway at a predetermined glide slope. After stabilizingat the desired flight conditions, the pilot attempted to hold the power constant and make minimal controlsurface inputs. As the airplane approached the runway and responded to ground effect, the pilotcontinued to hold the throttle constant and attempted to maintain a constant pitch attitude usinglongitudinal stick inputs. The maneuver was complete when either the airplane touched down or throttleadjustments were made.
Figure 3 shows a typical time history for a constant-alpha maneuver, including some of the maneuversetup and the data analysis time segment. The stage one compressor rate of rotation for engine number 1(
n1
) is shown to indicate periods of constant thrust. The value for engine number 1 is representative oftotal thrust, because the four throttles were normally adjusted in unison. The interval that starts whenthrust is held constant and ends with the onset of ground effect is referred to as the OGE period.
The constant-alpha approaches were attempted at a variety of OGE flight conditions; however, therange of flight conditions was constrained by the requirement to touchdown within acceptable verticaland horizontal speeds. Angle of attack was between 8.5° and 10.5° for all maneuvers. Weight changesrequired some variation in airspeed, but differences in canard position had a more significant impact onairspeed. Equivalent airspeed ranged from 170 to 181 knots with the canard extended and wasapproximately 200 knots with the canard retracted. Elevon position was also dependent on canard usage.The average elevon position was approximately 8° (trailing edge down) with the canard extended and –3°with the canard retracted. The pilot used the instrument landing system (ILS) glide slope indicator as anaid for setting up the initial glide slopes. Approaches were attempted at nominal flightpath anglesbetween –2° and –3°. Figure 4 shows a summary of the initial OGE flight conditions for the constant-alpha maneuver.
8
Figure 3. Typical time history for a constant-alpha maneuver.
9
Figure 4. Out-of-ground effect flight conditions for the constant-alpha maneuver.
Maneuvers were rejected from further analysis if winds exceeded 5 knots in any direction,inconsistencies in the measurements were observed, or thrust was not held constant below 1.5
h/b
. A totalof 19 constant-alpha approaches were attempted, of which 6 were used for analysis.
For the level pass flight maneuver, the pilot approached the runway on a glide slope between –2° and–2.6° and then transitioned to level flight. These passes were conducted during low-windspeed conditions(less than 5 knots in any direction) with the canard extended and the nose canted downward. A total of 7level pass flight maneuvers were conducted, from which 12 level segments were identified for analysis.To build confidence in the flight procedures, level passes were first conducted at 200 and 150 ft aboveground before working downward to as low as 24 ft (wing height) above ground. During the levelportions of the maneuver, the pilot attempted to minimize variations in thrust; however, throttleadjustments were generally necessary to adequately control the flightpath. Figure 5 shows a time historyfor a level pass maneuver. In this example, the airplane achieved two segments of level flight. The angleof attack and elevon positions during the second (and lowest) segment were approximately 5.2° and 5.9°,respectively. These conditions varied considerably from the OGE conditions for the constant-alphamaneuver (see figure 4).
The elevation of the runway varied considerably, and these maneuvers were actually flown parallel tothe runway surface. This requirement, along with general safety concerns of conducting a nonstandardmaneuver in close proximity to the ground, added considerable challenge to the level pass maneuver.
10
Figure 5. Typical time history for a level pass maneuver.
11
Data Analysis
Data analysis techniques for the constant-alpha approach were similar to the processes developed inreference 21. For each approach, a time interval was chosen during which the airplane was stabilized justbefore entering ground effect. Data from this OGE interval were averaged to determine OGE values forangle of attack, elevon position, flightpath angle, and thrust. The aerodynamic forces and moments actingon the airplane throughout the maneuver were then determined from the mass properties, thrust estimates,and measured body axis accelerations, using the following equations:
(1)
(2)
(3)
The pitching moments were calculated about the reference center of gravity. The forces were convertedto stability axis lift and drag coefficients using the following equations:
(4)
(5)
Variations in elevon position and angle of attack occurred during each maneuver and generally increasedas the airplane flew through ground effect. The effect of these variations had to be eliminated todetermine the direct influence of ground effect. The average values of elevon position and angle of attackduring the OGE portion of each maneuver were used as a reference from which the deviations could bemeasured. Aerodynamic derivatives were then used to extract the effects of trim changes that occurredduring the maneuvers, as follows:
(6)
(7)
Because the value of
C
m
was close to zero during the maneuvers, changes in angle of attack and elevonposition were the primary indications of ground effect. A table lookup function was used to correct thedrag coefficient data,
(8)
CX
Wax T–
Sq--------------------=
CZ
Waz
Sq----------=
Cm
Iyq I x Iz–( ) pr+
qSMAC-----------------------------------------=
CL C– Z α( )cos CX α( )sin+=
CD C– X α( )cos CZ α( )sin–=
CLrefCL CLα
α α OGE–( )– CLδeδe δeOGE–( )–=
CmrefCm Cmα
α α OGE–( )– Cmδeδe δeOGE–( )–=
CDrefCD CD α δ e , ( ) C D – α OGE δ e OGE , ( )[ ] –=
12
These terms, , , and are the aerodynamic coefficients of the airplane, referenced to the
OGE trim conditions.
The aerodynamic derivatives used in equations 6–8, such as and , and the drag function,
were obtained from reference 23 and were assumed to be largely independent of ground effect. Because
deviations from the reference conditions were typically small (< 3° in angle of attack, < 5° in elevon), this
assumption was considered acceptable. Because the engine throttles were not adjusted during the
constant-alpha approach, the incremental changes in the coefficients due to ground effect were not
significantly affected by errors in the estimation of thrust.
Aerodynamic forces and moments for each level pass maneuver were determined in a similar manner.During each segment of level flight, the total aerodynamic and thrust forces acting on the airplane weredetermined from accelerometer and mass property data. Because throttle settings changed during thelevel pass maneuver, the postflight estimate of engine thrust was a critical input to this analysis. Thisvariation in thrust was a significant disadvantage of the level pass maneuver compared to theconstant-alpha approach in which thrust was held constant.
For comparisons with other data sets, the data for each level flight segment were corrected to acommon reference trim condition (an angle of attack of 9° and an elevon position of 10°) using equations6, 7, and 8. Because angle of attack changed considerably between the various level pass maneuvers,these corrections were more significant for the level pass approach data than for the constant-alphaapproach data.
WIND TUNNEL TESTING
This section discusses the wind-tunnel testing of the Tu-144 airplane. The steady-state testsconducted by the Tupolev Design Bureau and both steady-state and dynamic tests conducted atNASA Langley are described in detail.
Tupolev Steady-State Testing
Wind-tunnel tests of the Tu-144 configuration in ground effect were conducted by the TupolevDesign Bureau during the initial development of the vehicle. A 4.8-percent scale model of theconfiguration was used, which included wing, body, vertical tail, nacelles, canards, and landing gear. Testconditions covered various heights above ground, flow angles, and control surface positions. The data aretabulated in reference 22. The wind-tunnel facility used for these tests did not incorporate a movingground plane or other boundary-layer removal system. Reynolds number was based on the modelmean aerodynamic chord (
MAC
), 3.67 ft.
NASA Langley Steady-State and Dynamic Testing
In October 1997, tests were conducted in the NASA Langley 14- by 22-ft Subsonic Wind Tunnel
(formerly the V/STOL Wind Tunnel and the Meter Subsonic Tunnel). This atmospheric tunnel has
CLrefCmref
CDref
CLαCLδe
4 106×
4 7×
13
a contraction ratio of 9 to 1, is powered by an 8000-hp drive system, and has a large test section capableof a maximum speed of 318 ft/sec. Details of the test facility can be found in reference 26. Although thetunnel can be operated in either a closed or open test section configuration, the dynamic ground-effecttests were conducted with a closed-flow system. A boundary-layer removal system, located just upstreamof the test section floor, was used.
A unique feature of the 14- by 22-ft Subsonic Wind Tunnel is that the model support hardware andfloor of the test section are combined into removable units known as carts. A special cart, called thedynamic ground effect (DGE) cart, was developed to enable testing while a model is moved within thepresence of ground effect. Figure 6 shows a diagram of the DGE cart. The large vertical support strut ishydraulically controlled to vary the model height above the cart (test section) floor. The strut has ahydraulic pitch drive that allows the model attitude to change during a dynamic plunge. The cart also hasa yaw drive to simulate the effects of sideslip, although it was not used in this test.
A computer system controls the operation of the hydraulic mechanisms. The system executes apreprogrammed model trajectory in which a sequence of vertical speed and pitch attitudes is prescribed.The range of vertical travel extends from 89 in. to approximately 5 in. depending on the model and modelpitch attitude. Model pitch limits range from –10° to +50°. Although the system is capable of a maximumvertical speed of 15 ft/sec, a maximum speed of 9 ft/sec was used in this test.
(a) Dynamic ground effect cart. (b) Tu-144 planform model and sting configuration.
Figure 6. NASA Langley 14- by 22-ft Subsonic Wind Tunnel model support arrangement.
14
The October 1997 test was the first operation of the DGE cart, and the primary objective was toconduct a shakedown of the new system and related instrumentation. A detailed summary of this test ispresented in reference 27.
One of several models tested during this developmental test was a planform representation of theTu-144 airplane (fig. 7). The model has a wingspan of 47.1 in. and is approximately 4.1 percent the sizeof the flight vehicle. The model is constructed of sheet aluminum, covered with wood, and nominallycontoured into a biconvex cross section with small radius leading edges. The elevon segments were set to10° trailing edge down for these tests. The reference center of gravity for the model is consistent with thefull-scale airplane (40 percent of the
MAC
).
Figure 7. Comparison of the Tu-144 airplane and Tu-144 wing planform wind-tunnel model.
The model was attached to the sting through a balance block that was enclosed in a fairing andmounted to the upper surface of the model. The model and balance block weighed approximately 60 lb.A 9° knuckle was used to provide more angle-of-attack capability near the ground plane.
Wind-tunnel instrumentation included a force and moment balance, accelerometers, and an opticaltracking system. Six model accelerometers were used in combination to provide three linear and threeangular accelerations of the model reference center. These data were used to remove inertial effects fromthe dynamic test data (discussed in the last paragraph of this section). Optical measurements were used todetermine the height of the model relative to the tunnel floor at two different model locations. These datawere used to define model height above the floor (ground plane), vertical speed, and pitch attituderelative to the tunnel axis system. Vertical speed data from differentiation of the position measurementswere confirmed by integration of sting-mounted accelerometer data.
15
All data in this study were obtained at a tunnel velocity of 267 ft/sec. Vector analysis was used todetermine the effective model airspeed, angle of attack, and flightpath angle from the tunnel speed andoptical tracking system data, as shown:
(9)
(10)
(11)
The wind-tunnel testing included a series of steady-state runs through a sweep of angles of attack.During these runs, the wing aerodynamic reference point was placed near the center of the tunnel at aheight of 6 ft (
h/b
= 1.53). This position was assumed to be essentially OGE. Data from these runs wereanalyzed using conventional methods. Tares were determined from measurements taken during wind-offconditions.
A second series of steady-state runs was conducted with the model positioned at lower heights abovethe ground plane (
h/b
= 1.019, 0.509, 0.382, and 0.256) and at angles of attack of 7°, 9°, and 11°. Theseruns provided steady-state ground-effect data for the wing planform model.
Prior to dynamic testing, the model was rapped with a rubber hammer to excite the natural vibrationmodes of the combined model and sting support system. These vibration data, taken with no airflow,were used to relate the model accelerations to inertial forces sensed at the balance.
The model support system was then used to plunge the model vertically toward the ground plane. Theeffective flightpath angles for these runs ranged from approximately –0.2° to –2.5°. The model pitchattitude was set as a function of vertical velocity for each run so that all runs were conducted at the sameeffective angle of attack (approximately 9.55°). All derived lift coefficients, therefore, were referenced tothis angle of attack using the OGE lift data for the wing planform.
Figure 8 shows typical time histories for two of the dynamic plunge tests. Data acquired from a lowvertical speed run (2.2 ft/sec) (fig. 8(a)) shows a brief acceleration interval followed by a significantperiod of constant-speed conditions. Large oscillations in the model speed and measured normal forceduring startup were the result of inertial forces induced by the rapid acceleration of the model andsubsequent “ringing” of the model and sting system. In this example, the inertial ringing ceased early inthe run. An attempt was made to remove the inertial forces from the measurements using the relationshipbetween accelerometer data and normal force data obtained during the wind-off vibrations. As shown infigure 8(a), this correction did not significantly improve the data during the initial acceleration but madea small improvement for heights between 0.3 and 1.3
h/b
. Figure 8(b) shows a time history for a higherspeed plunge (6.7 ft/sec). Because of the higher speed, most of the trajectory was affected by theacceleration and deceleration phases. The acceleration-based correction for inertial forces in
C
N
significantly improved the results, although it did not completely remove the inertial effects in the data.A test interval in which averaged conditions were constant was selected for these runs, althoughoscillations persisted throughout the run. Further details regarding the inertial correction process areprovided in reference 27.
V eff 2672
h2
+=
γeff sin1– h
V eff----------
=
αeff θ γeff–=
16
(a) Vertical speed approximately 2.2 ft/sec.
(b) Vertical speed approximately 6.7 ft/sec.
Figure 8. Time histories for dynamic plunge tests in the NASA Langley 14- by 22-ft Subsonic Wind Tunnel.
17
ANALYTICAL RESULTS
A constant-pressure, potential-flow, panel method (ref. 5) was used to predict ground-effectcharacteristics of the Tu-144 wing. A simple wing planform model was generated and used to determinethe OGE characteristics without the presence of a ground plane. No attempt was made to include thecorrect camber distribution or trailing-edge flap deflections in the model. The presence of the ground wasthen simulated by including an image of the isolated wing. The plane of symmetry between the twowings served as an effective ground plane. The aerodynamic forces in ground effect were determined byintegrating the panel pressures on the actual wing.
The resulting aerodynamic forces were then resolved into a lift component and drag component. Thedrag calculation includes only the lift-related or induced component of drag, because the panel code doesnot include viscous forces. In a similar manner, pitching moment about the aerodynamic reference pointwas determined by integrating the contributions from each panel.
Figure 9 shows results from the panel method. Because of the linear nature of potential flow theory,the computed forces were normalized by the corresponding OGE value. The normalized effects on liftand induced drag are identical.
Figure 9. Panel code normalized ground-effect predictions.
18
As suggested by analysis in reference 28, the variation of normalized ground-effect increments withnondimensional height above ground was found to fit a –1.5 power law relationship as shown:
(12)
A least-squares curve fit method was used to determine the coefficient, k, for the panel code lift anddrag data. As shown in figure 9, the resulting curve effectively represents the panel code data. As a result,the –1.5 power law relationship was also used in the interpretation of the experimental force dataacquired in this study. Pitching moment variations with height above ground are more complicated anddo not fit the power law relationship.
RESULTS AND DISCUSSION
Table 2 presents a summary of the nominal test configurations and conditions that are discussed inthis report.
Lift, drag, and pitching moment data from six constant-alpha maneuvers are shown as a function of non-dimensional height above ground (h/b) (fig. 10). As previously discussed, the data are shown in coeffi-cient form, referenced to the angle of attack and elevon position of the airplane immediately before enter-ing ground effect. The coefficient data are unaffected by proximity to the ground at heights above onespan (h/b = 1). Below this level, the influence of the ground on each coefficient is increasingly evident inall maneuvers.
Table 2. Summary of nominal test configurations and conditions.
Data Set α, deg δe, deg γ, deg
Flight (complete vehicle)
Dynamic (constant-alpha maneuver)
Canard extended 9 8 –2.1 to –3.0
Canard retracted 9.5 –2.5 –2.6 to –3.1
Steady-state (level pass maneuver) varies varies 0
Tupelov Wind Tunnel (complete vehicle)
Steady-state 6,10,15 0,5,10 0
NASA Langley Wind Tunnel (wing planform)
Dynamic 9.5 10 –.24 to –2.15
Steady-state 7,9,11 10 0
Analytical Data (wing planform)
Steady-state varies 0 0
∆CLGE
CLOGE
----------------- k hb---
1.5–=
19
(a) Flight 4.
Figure 10. Constant-alpha maneuver data.
20
(b) Flight 5.
Figure 10. Continued.
21
(c) Flight 12.
Figure 10. Continued.
22
(d) Flight 13.
Figure 10. Continued.
23
(e) Flight 14.
Figure 10. Continued.
24
(f) Flight 16.
Figure 10. Concluded.
25
Figure 10 also shows flightpath angle, angle of attack, and elevon position for each maneuver.Although these parameters were relatively constant before entering ground effect, the flightpath angle, γ,became less negative as the airplane approached the runway. This effect occurred while angle of attackremained constant or even decreased in some cases. This natural flaring of the flightpath is a distinctindication of increased lift due to ground effect. Although the constant-alpha maneuvers provided data ina dynamic ground-effect situation (γ ≠ 0), the rate of descent was not constant.
Because the pilot attempted to maintain a stabilized glide slope before entering ground effect,variations in the data obtained during the OGE portion of each maneuver (h/b > 1) indicate the quality ofthe force and moment measurements. The variations from the mean were calculated during the OGE timeintervals, and standard deviations in the lift, drag, and pitching moment coefficients were found to beapproximately 0.02, 0.003, and 0.002, respectively. These uncertainties are primarily attributed toatmospheric disturbances, sensor noise, and residual oscillatory dynamics (not included in the force andmoment calculations).
A least-squares method was used to fit the lift data to a –1.5 power law relationship with the heightabove ground, as discussed earlier. Although the resulting curves (also shown in figure 10) match thegeneral trend of the flight data, the fit is not as good as that of the analytical data in figure 9. Despite thiscase, the power law curve fit process is believed to provide a useful method for extracting data at specificaltitudes for correlation with other data sets.
The consistency of the lift and pitching moment data is better than that of the drag data, which isexpected, because the absolute magnitude of the drag increment caused by ground effect is much smallerthan the force increment in the lift axis. Similar results were observed in previous ground-effect flight teststudies (refs. 20, 21).
Figure 11 shows lift and pitching moment coefficients from the level pass maneuvers. As previouslydiscussed, the data have been adjusted to a common set of reference conditions (an angle of attack of 9°and an elevon position of 10°). Because thrust levels varied with each level pass maneuver, dragmeasurements were not reliable.
Figure 11. Level pass maneuver data.
26
Compared to the constant-alpha maneuver, the level pass maneuver has several deficiencies. First,seven level pass maneuvers were required to obtain the data in figure 11, whereas each constant-alphamaneuver provided an equivalent set of data. The level pass technique is clearly less efficient in terms offlight test operations. Second, the lift and pitching moment data from the level pass maneuvers are lessconsistent than the data from the constant-alpha approaches. Whereas the constant-alpha maneuver dataanalysis extracts ground effect from an incremental change in the forces and moments that occur duringthe same maneuver, the level pass maneuver data analysis requires correlation of data from multiplemaneuvers. This lack of reference is believed to have introduced errors that were not completelyaccounted for in the analysis process and is a fundamental disadvantage of the level pass test technique.
Figure 12 shows full-configuration, steady-state lift coefficient results at three elevon settings, each atthree angles of attack, from the Tupolev wind-tunnel test. When each set of ground-effect increments aredivided by the corresponding OGE lift coefficient, the normalized data sets collapse into a commontrend. Each set of data fits the –1.5 power law relationship reasonably well.
Similar results were found in the lift coefficient data from the wing planform model tests, which wereconducted in the NASA Langley wind tunnel under steady-state conditions (fig. 13). As shown, theconsistency of the wing planform data is very good.
(a) δe = 0°.
Figure 12. Full-configuration, steady-state wind-tunnel test data.
27
(b) δe = 5°.
(c) δe = 10°.
Figure 12. Concluded.
28
Figure 13. Steady-state wind-tunnel test data of the wing planform model, δe = 10°.
Figure 14 shows lift coefficient data, effective flightpath angle, and angle of attack from dynamicwind-tunnel testing of the wing planform model. Considerable scatter is present in both the trajectorydata and the lift coefficient data acquired at the beginning of each run (h/b ≈ 1.5). The scatter was causedby inertial forces and motions from the acceleration of the model. After this startup transient, the angle ofattack and flightpath angle were constant at the lower flightpath angles for the remainder of the two runs(figs. 14(a) and 14(b)). For runs at higher flightpath angles (higher vertical velocities), the angle of attackand flightpath angle decreased as the model decelerated at the end of the run. As previously discussed,the effective angle of attack during the constant-rate portion of each run was 9.55°, and all lift coefficientdata shown for this test have been referenced to an angle of attack of 9.55° for consistency.
To test the wind-tunnel measurements and analysis techniques, data derived from the dynamic runswere compared with steady-state measurements for the same model. For the dynamic tests, the planformmodel had a lift coefficient of 0.532 during the initial descent (above ground effect) at an effective angleof attack of 9.55°. These data compare well with the steady-state data.
For runs that simulated flightpath angles steeper than –1° (figs. 14(d), 14(e), and 14(f)), the entiretrajectories were at varying conditions. Although an increase in the lift coefficient caused by groundproximity is evident in every case, the quality of the data degraded to an unusable level as runs wereconducted at the higher descent rates.
29
(a) Nominal flightpath angle = –.24°.
(b) Nominal flightpath angle = –.48°.
Figure 14. Dynamic wind-tunnel test data of the wing planform model, δe=10°.
30
(c) Nominal flightpath angle = –.98°.
(d) Nominal flightpath angle = –1.49°.
Figure 14. Continued.
31
(e) Nominal flightpath angle = –2.17°.
(f) Nominal flightpath angle = –2.15 °.
Figure 14. Concluded.
32
To be able to compare the results from the various data sets discussed in this report, the liftincrements caused by ground effect were normalized to the corresponding OGE conditions. The powerlaw curve fit was applied to each data set and used to determine a normalized lift coefficient increment ata common height above ground (h/b = 0.3 was chosen, because it fell within the range of all the testtechniques). Figure 15 shows the resulting data as functions of flightpath angle, angle of attack, andelevon position.
Figure 15(a) shows no distinct variation in normalized lift coefficient increments as a function offlightpath angle, although the level pass value is larger than the other data. The normalized liftcoefficients from dynamic flight testing (γ < 0) are equivalent to the full-configuration, steady-statewind-tunnel data. Comparisons of the planform data sets in figure 15(a) also indicate no apparent trendresulting from a dynamic rate of descent (flightpath angle). With the exception of the level pass data, anyrelationship between lift coefficient ground effect and flightpath angle for this configuration is within themeasurement accuracy of these methods.
This finding conflicts with results from other studies (refs. 6–14) that have identified differences inground effect related to rate of descent. In some cases, particularly for highly swept configurations, ratiosof steady-state to dynamic ground-effect coefficients as large as 2 have been reported. Such a large ratiois not evident in the Tu-144 data sets.
Figure 15(a) also indicates that ground effect for this vehicle is not a strong function of otherparameters. For example, the full-configuration wind-tunnel test data, which had no engine exhaust flowsimulation and were obtained at low Reynolds numbers, compare reasonably well to the full-scale flighttest data. Furthermore, results from the simple, uncambered wing planform model compare reasonablywell to those of the flight vehicle. These conclusions imply that complex, dynamic wind-tunnel testingmight not be required to characterize ground effect for this particular configuration. Further research isrecommended to identify configuration or flight condition parameters that indicate the presence ofdynamic effects. Additional experimentation under controlled conditions with systematic variations inconfiguration parameters (such as leading-edge vortex strength, and so forth) and trajectory might lead toa satisfactory understanding of the DGE problem. The DGE wind-tunnel facility would be ideal for thispurpose if the quality of the data obtained at comparatively higher dynamic rates could be improved.
Figure 15(b) shows the normalized lift coefficient data as a function of angle of attack. Allconstant-alpha approach flight data were obtained at angles of attack near 9°. Level pass maneuvers wereconducted at angles of attack that ranged from 5 to 8°; however, results from each maneuver wereadjusted to a reference angle of attack of 9° and an elevon position of 10° using free-stream derivatives,as previously discussed. These corrections (which are large compared to the corrections made to theconstant-alpha method) further question the accuracy of the level pass maneuver data. In the wingplanform results, no similar variation in angle of attack is observed.
Variations in the normalized lift coefficient as a function of elevon position are relatively small forthe full-configuration tests (fig. 15(c)). All wing planform data were obtained at an elevon position of10°, and the linear panel code data were independent of elevon position.
33
(a) Varying flightpath angle.
9° ≤ α ≤ 10°.
9° ≤ δe ≤ 10°.
(b) Varying angle of attack and flightpath angle.
9° ≤ δe ≤ 10°.
Figure 15. Comparison of normalized ground-effect increments, h/b=0.3.
34
(c) Varying elevon position and flightpath angle.
9° ≤ α ≤ 10°.
Figure 15. Concluded.
CONCLUDING REMARKS
Ground-effect characteristics of the Tu-144 supersonic transport airplane have been obtained fromboth dynamic and steady-state flight test maneuvers, and from a steady-state, full-configurationwind-tunnel test. A simple model of the Tu-144 wing planform was also tested under both dynamic andsteady-state conditions using a developmental wind-tunnel model support system. Panel code predictionsof ground effect for the wing planform were also obtained. Ground-effect increments from the variousstudies were compared by normalizing the coefficient data to out-of-ground effect (OGE) conditions andfitting the results to a power law trend.
Aerodynamic Findings
Results from the panel code analysis and steady-state wind-tunnel testing of both the full-configuration and wing planform models were found to be highly consistent and fit the power lawrelationship very well. Results from the dynamic flight tests, obtained during constant-alpha approaches,exhibited more scatter and some oscillatory trends, but these results compared favorably with the datafrom the steady-state, full-configuration wind-tunnel tests. Data from the steady-state flight tests,obtained during the level pass maneuvers, indicated slightly higher values of ground effect in the lift axis,but the data quality was relatively poor.
35
Data from the dynamic wind-tunnel tests were highly consistent for runs at low descent rates. TheOGE data from these runs successfully matched the steady-state wind-tunnel data for the same model,further validating the dynamic test technique. Unfortunately, the effective flightpath angles for these runswere lower than those of the typical airplane landing flightpath. Data quality at higher descent rates wasdiminished by relatively large inertial forces, which complicated the balance measurements and limitedthe consistency of the model trajectory. Despite these limitations, the dynamic wind-tunnel testsindicated no significant variation in ground effect as a function of effective flightpath angle for thisconfiguration. This finding conflicts with findings from previous studies of low-aspect-ratio, swept-wingconfigurations, which showed large differences between dynamic and static ground effect. Furtherresearch is recommended to identify controlling parameters that are responsible for these differences.
Test Technique Findings
The current dynamic wind-tunnel test data set presented in this report had several limitations, but thistest technique might be highly valuable in future ground-effect research. In particular, the current datawere obtained during pathfinder testing of the newly developed dynamic ground effect (DGE) modelsupport system. Future improvements based on the current experience might lead to a capable system interms of both operational effectiveness and test envelope. This type of system might be required to obtainparametric data to determine the controlling aspects of DGEs.
Although flight testing offers the only opportunity to validate ground-effect predictionmethodologies, this type of measurement remains a challenging undertaking. The optimal flight test datain this study were obtained during constant-alpha approaches in which the results could be extracted fromincremental changes from OGE reference conditions. Although 19 constant-alpha approaches wereattempted, only 6 were acceptable for analysis. In future projects, more flight test time should be devotedto practicing and repeating the maneuvers to ensure a sufficient set of usable data. The level passmaneuvers required substantially more flight test time and were of limited value, primarily because ofuncertainties in the determination of thrust variations between maneuvers.
REFERENCES
1. Nelson, C.P., “Effects Of Wing Planform On HSCT Off-Design Aerodynamics,” 10th AppliedAerodynamics Conference, AIAA-92-2629, June 1992.
2. Kemp, William B., Vernard E. Lockwood, and W. Pelham Phillips, Ground Effects Related toLanding of Airplanes With Low-Aspect-Ratio Wings, NASA TN D-3583, Oct. 1966.
3. Snyder, C. Thomas, Fred J. Drinkwater III, and A. David Jones, A Piloted Simulator Investigationof Ground Effect on the Landing Maneuver of a Large, Tailless, Delta-Wing Airplane, NASATN D-6046, Oct. 1970.
4. Anderson, John D., Jr., Fundamentals of Aerodynamics, McGraw-Hill Book Company, 1984.
5. Hedman, Sven G., Vortex Lattice Method for Calculation of Quasi Steady State Loadings on ThinElastic Wings in Subsonic Flow, The Aeronautical Research Institute of Sweden, Report 105, 1966.
36
6. Chen, Yen-Sen and William G. Schweikhard, “Dynamic Ground Effects on a Two-Dimensional FlatPlate,” Journal of Aircraft, vol. 22, no. 7, July 1985, pp. 638–640.
7. Nuhait, A.O., “Unsteady Ground Effects on Aerodynamic Coefficients of Finite Wings withCamber,” Journal of Aircraft, vol. 32, no. 1, Jan.–Feb. 1995, pp. 186–192.
8. Nuhait, A.O. and M. F. Zedan, “Stability Derivatives of a Flapped Plate in Unsteady Ground Effect,”Journal of Aircraft, vol. 32, no. 1, Jan.–Feb. 1995, pp. 124–129.
9. Chang, Ray Chung and Vincent U. Muirhead, “Effect of Sink Rate on Ground Effect of Low-Aspect-Ratio Wings,” Journal of Aircraft, vol. 24, no. 3, March 1986, pp. 176–180.
10. Lee, Pai-Hung, C. Edward Lan, and Vincent U. Muirhead, An Experimental Investigation ofDynamic Ground Effect, NASA CR-4105, Dec. 1987.
11. Chang, Ray Chung, An Experimental Investigation of Dynamic Ground Effect, PhD Dissertation,University of Kansas, 1985.
12. Kemmerly, G.T., J.W. Paulson, Jr., and M. Compton, “Exploratory Evaluation of Moving-ModelTechnique for Measurement of Dynamic Ground Effects,” Journal of Aircraft, vol. 25, no. 6,June 1988, pp. 557–562.
13. Kemmerly, Guy T. and John W. Paulson, Jr., Investigation of a Moving-Model Technique forMeasuring Ground Effects, NASA TM-4080, Jan. 1989.
14. Paulson, John W., Jr., Guy T. Kemmerly, and William P. Gilbert, “Dynamic Ground Effects,”AGARD Fluid Dynamics Panel Symposium on Aerodynamics of Combat Aircraft Controls and ofGround Effects, AGARD CP-465, April 1990, pp. 21-1–21-12.
15. Rolls, L. Stewart and David G. Koenig, Flight-Measured Ground Effect on a Low-Aspect-RatioOgee Wing Including a Comparison with Wind-Tunnel Results, NACA TN-D-3431, 1966.
16. Rolls, L. Stewart, C. Thomas Snyder, and William G Schweikhard, “Flight Studies of GroundEffects on Airplanes with Low-Aspect-Ratio Wings,” Conference on Aircraft Aerodynamics,NASA SP-124, May 1966, pp. 285–295.
17. Baker, Paul A., William G. Schweikhard, and William R. Young, Flight Evaluation of Ground Effecton Several Low-Aspect-Ratio Airplanes, NASA TN D-6053, 1970.
18. Schweikhard, William, “A Method for In-Flight Measurement of Ground Effect on Fixed-WingAircraft,” Journal of Aircraft, vol. 4, no. 2, Mar.–Apr. 1967, pp. 101–104.
19. Curry, Robert E., Bryan J. Moulton, and John Kresse, An In-Flight Investigation of Ground Effect ona Forward-Swept Wing Airplane, NASA TM-101708, Sept. 1989.
20. Corda, Stephen, Mark T. Stephenson, Frank W. Burcham, and Robert E. Curry, Dynamic GroundEffects Flight Test of an F-15 Aircraft, NASA TM-4604, Sept. 1994.
37
21. Curry, Robert E., Dynamic Ground Effect for a Cranked Arrow Wing Airplane, NASA TM-4799,Aug. 1997.
22. Task 39 Tu-144LL Follow-on Program: Final Report, HSR-AT Contract No. NAS1-20220,Submitted by The Boeing Company to NASA Langley Research Center, June 1999.
23. Cox, Timothy H. and Alisa Marshall, Longitudinal Handling Qualities of the Tu-144LL Airplaneand Comparisons With Other Large, Supersonic Aircraft, NASA TM-2000-209020, May 2000.
24. Rivers, Robert A., E. Bruce Jackson, C. Gordon Fullerton, Timothy H. Cox, and Norman H. Princen,A Qualitative Piloted Evaluation of the Tupolev Tu-144 Supersonic Transport, NASA TM-2000-209850, Feb. 2000.
25. Ashtech™ Precise Differential GPS Navigation (PNAV) Trajectory Software User’s Guide,Software Version 2.1.00-T, Document Number 600200, Revision B, May 1994.
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27. Graves, Sharon S., Investigation of a Technique for Measuring Dynamic Ground Effect in aSubsonic Wind Tunnel, NASA CR-1999-209544, Aug. 1999.
28. Hoerner, Sighard F. and Henry V. Borst, Fluid-Dynamic Lift: Practical Information on Aerodynamicand Hydrodynamic Lift, Second edition, Published by Mrs. Liselotte A. Hoerner, 1975.
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NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)Prescribed by ANSI Std. Z39-18298-102
Ground-Effect Characteristics of the Tu-144 Supersonic TransportAirplane
710-55-04-E8-RR-00-000
Robert E. Curry and Lewis R. Owens
NASA Dryden Flight Research CenterP.O. Box 273Edwards, California 93523-0273
H-2510
National Aeronautics and Space AdministrationWashington, DC 20546-0001 NASA/TM-2003-212035
Ground-effect characteristics of the Tu-144 supersonic transport airplane have been obtained from flight andground-based experiments to improve understanding of ground-effect phenomena for this class of vehicle. Theflight test program included both dynamic measurements obtained during descending flight maneuvers andsteady-state measurements obtained during level pass maneuvers over the runway. Both dynamic and steady-state wind-tunnel test data have been acquired for a simple planform model of the Tu-144 using a developmentalmodel support system in the NASA Langley Research Center 14- by 22-ft Subsonic Wind Tunnel. Data fromsteady-state, full-configuration wind-tunnel tests of the Tu-144 are also presented. Results from theexperimental methods are compared with results from simple computational methods (panel theory). A powerlaw relationship has been shown to effectively fit the variation of lift with height above ground for all data sets.The combined data sets have been used to evaluate the test techniques and assess the sensitivity of ground effectto various parameters. Configuration details such as the fuselage, landing gear, canards, and engine flows havehad little effect on the correlation between the various data sets. No distinct trend has been identified as afunction of either flightpath angle or rate of descent.
Aerodynamics, Flight test, Ground effect, Supersonic transport, Tu-144 42
Unclassified Unclassified Unclassified Unlimited
October 2003 Technical Memorandum
Unclassified—UnlimitedSubject Category 02
This report is available at http://www.dfrc.nasa.gov/DTRS/